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Why does the A320 use the rudder for lateral control in mechanical law?
When the Airbus A320’s flight control system is operating in mechanical law (the simplest and lowest-tech of the A320’s flight control laws, where deflections of the pilots’ joysticks are transmitted directly to the control surfaces via ye olde hydraulickes, with absolutely no envelope protections at all; mechanical law is the A320’s only completely non-fly-by-wire flight control law, and is generally only used in the event of an utter and total failure of the aircraft’s flight control computers), the aircraft’s ailerons and spoilerons are inoperative, and lateral control is accomplished solely by means of the rudder.
However, the rudder is a horrible lateral control surface; its intended purpose is for directional control, controlling yaw, not roll.  Although an aircraft’s rudder can, in a pinch, be used to roll the aircraft in whichever direction the rudder is deflected, it does this indirectly, by inducing sideslip.  The rudder, being quite close to the aircraft’s centerline, produces only a small direct rolling moment - which, to make matters worse, for aircraft with the rudder mounted above the aircraft’s center of mass (e.g., essentially all commercial airliners, the A320 included), is in the wrong direction.  Thus, if one wants to turn right, and applies right rudder to that effect, the aircraft will initially roll slightly to the left, then stop, then, as the nose yaws to the right and the sideslip angle grows, finally roll suddenly and violently to the right.  Once you’re actually in a turn, the situation is improved not at all, as now you have to religiously avoid applying any further rudder in the direction of the turn, lest you start skidding (turning with an excess of yaw, resulting in the aircraft’s nose pointing to the inside of the turn), which is an excellent, time-tested way of entering an unrecoverable spin and making your insurance company give your family lots of money.
If one had to choose just one set of control surfaces to use when turning, one would think that the optimal choice would be the lateral controls (ailerons and spoilerons), not the directional controls (rudder); that way, the required turning technique would be essentially unchanged from during normal flight, as modern aircraft do not use the rudder at all under normal circumstances (advances in aileron design1 having freed newer designs from the severe adverse yaw that bedeviled the Wrights and was the rudder’s original raison d’etre).2  Indeed, the A320’s stateside counterpart, the Boeing 737, does exactly this; when operating in manual reversion (the closest 737 equivalent to the A320’s mechanical law, given that the A320 inexplicably has no manual-reversion capability despite being small enough that it should have been a breeze to include), the 737’s rudder is (nominally) inoperative,3 and the aircraft is turned using only the ailerons and spoilerons.
Given the immense superiority of the ailerons and spoilerons over the rudder for lateral control, why did the A320’s designers go with the latter?

1: Consisting of various ways of artificially increasing the drag experienced by the falling wing in order to balance out the increase in induced drag on the rising wing resulting from its downturned aileron.  The major way of doing this is simply by adding spoilerons to the system, which has the double benefit of eliminating adverse yaw (or even turning it into proverse yaw - that is, yaw in the direction of the turn, rather than out of it) and considerably increasing roll authority (especially at low speeds - and, hence, high angles of attack - where the maximum downwards aileron deflection is severely restricted to avoid stalling the rising wing, which would produce a rolling moment in the wrong direction, somewhat hampering one’s efforts to turn the plane in the direction intended); other methods include deflecting the upturned aileron on the falling wing much further from its faired position than the downturned aileron on the rising wing (thus balancing out the increase in induced drag on the downturned aileron with an equal or greater increase in pressure drag - and, at the transonic cruising speeds of modern jetliners, potentially wave drag as well - on the upturned aileron) and building the ailerons such that, when deflected upwards, part of the hinge end of the aileron protrudes below the wing (where it produces extra drag on the now-falling wing to balance out the extra drag on the now-rising wing).
2: In modern aircraft, the rudder is only used when it is necessary, for some reason, to generate a large sideslip angle, with these cases falling into three general categories:

The rudder’s design case (the scenario that places the greatest demand on the rudder, and, thus, determines the minimum control authority - and, thus, indirectly, the size - required of the rudder) is an engine failure at low speed, with the absolute worst case being a sudden, total failure of the furthest-outboard engine during takeoff, immediately after V1.  At these low speeds, the rudder has to be large - especially on wing-engined aircraft - to counter the large yawing moment from a failed engine, which is why the rudder has enough control authority to overpower the lateral controls at low speed, which is why all aircraft have a crossover airspeed.
The most common use of the rudder, in contrast, is during crosswind landings, where the aircraft, in order to keep its nose pointed into the relative wind during the descent and flare, has to approach the runway with a large crab angle (i.e., with the aircraft’s longitudinal axis forming a large angle with that of the runway, rather than the two being parallel, or, ideally, coinciding), and then, upon mainwheel touchdown, must immediately (within a second or two) yaw to the runway heading (which requires a very high yaw rate, especially with a high crosswind component and resultant very large crab angle, and, thus - as the tiller can’t be used for steering until the nosewheel touches down, and, at any rate, would be largely ineffective, except at destroying the nosewheel tyres, at the high speeds of a touching-down jetliner - a sudden, large rudder input) in order to avoid running off the side of the runway.
The third main rudder-use scenario is if the aircraft has to turn suddenly when moving at speed on the ground (for instance, to avoid colliding with a baggagecart/deer/other plane/very lost driver/inattentive groundcrewperson/tiger griffin/tumbleweed that’s somehow found its way onto the runway you’re using to take off and/or land).  In this case, the aircraft cannot roll, because the ground is in the way, so the rudder (along with differential braking, and, at lower speeds, some tiller input) is used to turn the aircraft.

3: In actuality, a small amount of rudder deflection is available in manual-reversion flight if enough force is applied to the rudder pedals, due to the specifics of the design of the 737’s rudder control system.

You might like to consider making your questions a little shorter and focus on the key points. A long, detailed question suggests - rightly or wrongly - that a long, detailed answer may be needed. That tends to turn off people who could answer, and it doesn't fit the StackExchange philosophy of specific, answerable questions.

You normally never touch the rudder pedals in a jet once airborne unless an\ engine fails or you are landing.  But if you've ever tried applying rudder-only in a swept wing jet, you quickly discover you can get pretty snappy roll rates with very small amounts of yaw.

While long, it is a well written question, "which is an excellent, time-tested way of entering an unrecoverable spin and making your insurance company give your family lots of money."

The joysticks don't have any cables, etc. attached to them. In mechanical back-up, where electrical systems are lost, there's just no way to have them control the ailerons.

What a great, well-asked question

You've been heard: Backup Control Module

OK, what you call "mechanical law" isn't a "law", it's mechanical signalling. So it needs steel cables from somewhere in the cockpit to whatever chosen flight controls.
The A320 already uses steel cables for the Rudder - that's how it's normally signalled. So there's no extra work involved in doing this. For pitch control, the Stab Trim Wheel is never used in normal flight, but has steel cables going to the Stab, so this is used in mechanical signalling mode.
The aircraft is perfectly flyable (if you broaden the definition of 'flyable' a little) in mechanical signalling mode, but lots of things have to fail to get you down to that level.

very interesting answer; may be you may be helpful for the following:I discovered lately that the mechanical backup of the B777 apparently  doesn’t use the rudder but only the THS and 2 symmetrical spoilers(N°4 and N°11), however to loose totally the electrical power and to exhaust the batteries I am obliged to imagine the loss of an engine, so my worries how to fly the aircraft without the rudder in this case?Is it possible the THS and 2 spoilers to be sufficient with a modulated power of the remaining engine? If you have a solution please answer the question I have posted. Kind Regards

However, the rudder is a horrible lateral control surface; its intended purpose is for directional control, controlling yaw, not roll.

Yes indeed. But ultimately, in order to direct an aeroplane to the desired location, we need to control three degrees-of-freedom: fwd/aft, left/right, up/down.

fwd/aft control is achieved via the engines, or by trading in potential energy for kinetic energy during a glide.
left/right control is via rolling and/or yawing the aeroplane. Rolling is only preferred because it is much more comfortable for anyone, however when the chips are down in rare situations, comfort is not the first priority.
up/down via the elevators or horizontal tail.

Mechanical back-up in the A320 is for being able to control the aeroplane when all the flight control computers are down and out. And the rudder provides the necessary back-up control, more suitable for landing in a side wind than the ailerons do.

"And the rudder provides the necessary back-up control, more suitable for landing in a side wind than the ailerons do."- I wonder if this is true.  Would be interesting to play with a flight simulator and find out, w/ a realistic flight model of an airliner. Consider that--even in zero wind-- if you make a have to make a last-second roll correction with the rudder to stop yourself from touching down in a banked attitude and dragging a wingtip, this may involve so much yaw and slip that you may touch down with more sideload on the gear than you would in a no-rudder landing in a strong xwind.

I know for sure if I were landing, say, a Cessna 172 I would much rather have aileron and not rudder rather than vice versa.  The roll authority is much better in the former case, and so what if touchdown is highly crabbed in a sidewind with no ability to do either a cross-countrolled-sideslip or a last-second "de-crab" maneuver?  The landing gear can probably take it.  On the other hand I've flown a small number sailplanes w/ small vert fins that are literally uncontrollable w/ pilot's feet off rudder pedals.  In such cases it would be better to have only the rudder than only the ailerons.

But if the rudder on A320 is assumed locked near centered rather than free-floating and pilot is controlling roll with only ailerons or only spoilerons, I would bet the resulting flight characteristics would be far superior to the case where the ailerons are locked near center and the pilot is trying to accomplish roll control with only the rudder.  I may be wrong.  Ample dihedral and resulting unfavorable coupling between adverse yaw and roll is one reason the sailplane w/ small fins was uncontrollable w/ free-floating rudder, and airliners do have significant dihedral effect due to sweep.


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Why do newer 737s use two different styles of split winglets?
All new 737s use split-tip winglets, which are more efficient than the simpler blended winglets previously used.
However, for no clear reason, they use two different types of split-tip winglet:

New 737 Next Generations (the 737-700/-800/-900)1 use split-scimitar winglets, which have a main body shaped like a blended winglet, but with its upper tip hooked backwards, and also have an additional scimitar-shaped fin projecting outwards and backwards from the upbend in the winglet.

(Image by Mnts at Wikimedia Commons.)

The 737 MAX series uses the eponymous MAX winglets, which have a simpler shape than the split-scimitar winglet, looking essentially like if someone decided to put a winglet on the bottom of the wingtip in addition to the one at the top.

(Image by Aka The Beav at Flickr, via Helmy oved at Wikimedia Commons, modified by Altair78 at Wikimedia Commons.)
I don’t get it - why go to the trouble of producing two slightly different styles of winglet alongside each other for two mostly-similar families of the same aircraft?  Why not just go with the more efficient style of the two (be it the split-scimitar winglet or the MAX winglet), and use that on both the Next Generation and the MAX?

1: The 737-600 also belongs to the Next Generation family, but it left production long before the Next Generations switched from the blended to the split-scimitar winglet.

@ymb1 The 737 FRACTAL will have an infinite tree of smaller and smaller winglets. This will be so efficient, it will land with more fuel in the tanks than when it took off.

@DavidRicherby: To get regulatory approval for that, the 737 FRACTAL will dump fuel before landing so that pilots will not have to be trained to deal with the heavier landing off weights. Of course, there will be no training so fuel may be dumped on the same school under the flight path multiple times per day, and the US will be the last to ban the plane.

If I had to guess, the two different winglets have been added probably because of the certification process of the retrofit NG winglets. The scimitars still look a lot like the older winglets, which means they probably could install the changed winglet tip and the added lower winglet without having to change much of the outer wing structure and without expensive costs for certifying a totally different design. Even the position of the lights changed on the MAX - that all needs to be certified and they probably wanted to save money at the time they added the scimitars.

related:Is wingtip design mature?

The 737NG was originally introduced with with no winglets. A company called Aviation Partners worked with Boeing to develop the "blended" winglet, originally for the NG-based BBJ (Boeing Business Jet). Customers then had them installed after delivery for a while before Boeing worked their own version into the production line. There is a similar case now with the split scimitar, where Boeing is delivering the 737NG with the "blended" winglet and customers are having it replaced with the split scimitar from Aviation Partners Boeing sometime after delivery if they choose.
One reason for developing a new winglet on the MAX is that it helps differentiate it from the NG and has become part of its "brand."
There also may be a consideration for cost from certifying the new winglets. As the split scimitar and MAX designs are so similar overall, there may not be enough benefit of the MAX design to justify certifying it for the NG, where they needed to do extensive flight testing on the MAX anyway. The split scimitar is also somewhat a modified blended winglet, which may have made certification easier.

Like I said, "styling".

Since a winglet is basically a sail generating thrust (lift with a modest forward-tilted vector) from the circulating flow around the tip, extracting beneficial energy from the circulation (the thrust part) and producing an "outwash" that opposes and weakens the vortice flow in the process, it's just adding another sail beside the first one to extract more of the available energy from the flow.  From the NASA Dryden page:

Winglets, which are airfoils operating just like a sailboat tacking upwind, produce a forward thrust inside the circulation field of the vortices and reduce their strength. Weaker vortices mean less drag at the wingtips and lift is restored. Improved wing efficiency translates to more payload, reduced fuel consumption, and a longer cruising range that can allow an air carrier to expand routes and destinations.
To produce as much forward thrust as possible, the winglet's airfoil is designed with the same attention as the airfoil of the wings themselves. Performance improvements generated by winglets, however, depend on factors such as the basic design of the aircraft, engine efficiency, and even the weather in which an aircraft is operating.

In the case of the MAX, you can see that the lower winglet is at less than a 45 degree angle, so it's producing mostly vertical lift with a slight forward component because as you can also see its incidence is somewhat nose down relative to the main wing to optimize its AOA in the tip circulation.  It's kind of a half winglet, half tip extension.
Different engineering groups will do studies of different configurations and will say, "hey, if we do this, it'll be some little bit more efficient than if we do that, based on our particular analysis".
If you put different groups of engineers together to attack the problem, they are all going to come out with variations on what they think is the ideal configuration.  And so you see seemingly endless permutations that are, really, mostly nibbling at the margins of the major benefit that was achieved when you put one there in the first place.
Plus there is probably a bit of "styling" going on as well.  Just to be different.

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Why are the Convair 880's yokes shaped differently from each other?
Looking at this picture of a Convair 880's flight deck...

(Image originally by the San Diego Air & Space Museum at Flickr, via Wikimedia Commons)
...the captain's and first officer's control yokes have quite different shapes; the captain's yoke is round, like that of an automobile or a 1930s propliner, while the first officer's yoke has the familiar W-shaped yoke, complete with buttons on the yoke's horns.
Why the difference between the two yokes?  Is it to optimise the captain's yoke for the duties of a pilot flying, and the first officer's yoke for the duties of a pilot monitoring?  Or something else?

My guess is that this is a relatively odd situation, where one yoke had already been replaced by a newer one (on the right), whereas the other still is the original.
If you look at the two images below (from Airliners.net), both yokes are the same.

It seems as if the right one has some more controls on the yoke (e.g. for radio settings, trim controls, or something else) and is simply an upgrade to the older model. If the old one still worked, I can imagine an airline waiting for a regular service inspection or overhaul before replacing it, rather than removing parts that are still fit to fly.

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What makes it safe for the DC-8 to use reverse thrust in flight?
For most jets, the use of reverse thrust to slow down in flight, where even possible, is extremely dangerous, and strictly prohibited.
Not so on the DC-8, which is certified for the safe in-flight use of reverse thrust on the two inboard engines throughout its entire flight envelope, and which, in fact, uses in-flight reversing as a standard part of a normal descent - this despite the severe disruption in-flight reversing causes to the airflow over the wing.
What is it about the DC-8 that makes it capable of safely using reverse thrust in flight, unlike other jets?

The IL-62 can do that, too; I think NASA modified a 727 for the Shuttle training to do the same. Trident allegedly could do it, as well, and on the military side, I believe the C-17 can use reverse in flight, too. There might be more!

@CptReynolds -- not sure about T-tail/rear-engine designs, but my understanding is the C-17 is able to deploy its reversers in flight, something used to allow for steep tactical penetration descents.  Maybe someone with a C-17 flight manual handy could illuminate us sometime?

All that is needed to allow this is a strengthened wing structure that will safely absorb the forces created by using thrust reversers in flight. Most aircraft don't need that because they have effective speed brakes.

Concorde also used reverse thrust on 2 engines when performing an emergency descent due to pressurization failure to get to save altitudes fast enough.

No kidding, losing cabin pressure at 50-60k feet would definitely be a hair-raising event. I wonder if it ever happened on a commercial flight.

It is still considered somewhat dangerous to deploy the thrust reversers on a DC-8, but because the pilot can only do so on the inboard engines, any extreme vibration or buffeting is transmitted closer to the wing root, which is safer than doing so on the outboard engines. Farther outboard the force of the buffeting might damage the wing's mechanisms or structure. Still, it was rarely used because even when the pilots warned the passengers, it still scared the daylights out of them.
The DC-8 also has ground spoilers, but they were not meant to be used in flight. The one time they were used in flight, they caused one of the first fatal accidents on the type. The lack of spoilers/speedbrakes that could be used in flight made it very difficult to do steep approaches without excessively increasing airspeed, hence the ability to deploy the inboard thrust reversers in flight to compensate.

Juan is correct, inflight reversing was a bit noisy; however, it was a MEL requirement to have the inboard reversers operational to operate above FL350. The procedure was as follows:
For a steep or emergency decent all throttles are retarded to idle and the inboard thrust brake levers are put into "Reverse Detent Thrust" and when the "Eng Thrust Brake" indicating lights come on the throttles are moved to the reverse power stop. The outboard engines will be at flight idle as reverse in engines 1 & 4 is not possible until the landing gear is extended. Additionally, the use of reverse thrust was to be discontinued below 190 knots.
Sorry Juan, no speed brakes on the DC-8, but yes on the DC-9. The spoilers on the DC-8 were used to augment flight control (three outboard on each side) but only when the gear was down. This made handling the aircraft in gusty weather much easier giving a quick roll response. On landing the spoilers(two inboard each side and three outboard) would deploy when armed by a signal from rear wheel spin-up or, as a back-up, by nose gear compression. The crew could also deploy the spoilers manually by pulling the spoiler lever back and up.

Does this also apply to the variants with more modern CFM56 engines?

"Additionally, the use of reverse thrust was to be discontinued below 190K." - I'm guessing you mean 19k or FL190? I don't think DC-8s usually go up to 190k.

No, you guessed incorrectly! This is a DC-8 NOT a DC-3. The VNE (KIAS) at 10,000 ft is 352 Knots. MMO above FL29.3 is .88M. We normally tried to discontinue reversing on the ground below 80K; outboard engines first followed by the inboard engines to prevent compressor stall.

@Loren, the problem is that 190K is a bit unclear and Terran clearly understood it as altitude. Do you mean speed 190 knots?

Of course I mean speed, 190K translates in any reference or manual as (knots of indicated airspeed). Altitude is expressed in feet, meters or as a FL (flight level).

Knots is usually abbreviated as kt, not k. :) I will correct the mention of the speed brakes.

No Juan, wrong again! The flight manuals use only the "K" for speeds and the correct abbreviation is "Kts", not Kt. But, correct, no speed brakes or inflight spoilers on the DC-8 to reduce speed, only in-flight reverse.


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Why do pilots typically start out in piston-engined aircraft, rather than jets or turboprops?
Essentially all pilots begin flying with propeller aircraft powered by piston engines, rather than in turboprops or light jets, even though:

Operating a piston engine requires adjusting the mixture ratio, carburetor heating, ignition timing, and other such voodoo, whereas, with a turbine engine, you basically just turn on the ignition, press the START button, and you’re in business.
Jets and turboprops are orders of magnitude more reliable than piston engines, despite the latter’s decades-long head start in development; destructive failures of aircraft piston engines are much less common nowadays than in the 1940s and 1950s, when they were practically a once-a-flight occurrence, but piston engines still come up short compared to jets, most pilots of which go their entire careers without experiencing a single engine failure of any kind whatsoever.
And when a turbine engine does fail, it does so loudly and unmistakeably - none of those deadly invisible engine failures that killed off the market for piston-engined twinmotors, along with many thousands of their pilots.
In all but the coldest climates, jets and turboprops use Jet A (in the United States) or Jet A-1 (the rest of the world), two very similar kerosenes that are far safer to handle than the avgas burned in most piston engines (Jet A and A-1 have much lower vapour pressures, much higher flashpoints, and much slower flame spread rates than avgas, and, unlike avgas, are not full of tetraethyllead), and more widely available to boot; even in those extremely cold climates where the more enthusiastically flammable Jet B (a gasoline-kerosene blend) has to be used instead, it’s still safer than avgas.
Also, a turbine engine will run happily on avgas if the ground crew refuels it from the wrong container (jets are rightly famous for running on anything that flows and burns, as long as it doesn’t produce an exhaust containing refractory oxide particles), whereas accidentally fuelling a piston engine with jet fuel will cause it to blow its insides to bits just long enough after takeoff for you to now be over a residential subdivision or something similarly inconvenient for a forced landing.
Propellers come with their own complications; the propwash coming off the propeller and hitting the wings and tail creates nasty asymmetrical rolling and yawing tendencies, the pitch of most propellers has to be carefully managed (at least during ground operations, and sometimes in flight as well), and an even slightly nose-low landing, a gust from the wrong direction at the wrong moment, or a bump or pothole in the surface can easily cause an engine-destroying propstrike.

Given all this, why do pilots generally make their debut in piston-propeller aircraft, instead of in light jets or turboprops?

That question asks more in general; this one raises several specific points that would seem to favour learning to fly in jets, or at least turboprops, rather than in piston aircraft.

The only new point I see here is jet fuel vs avgas, and why nobody makes recipe engines that run on jet fuel is a good question but only marginally related to safety of trainers.

I marked this a dupe because even if all of your individual arguments are valid, there are other points that are even more important and those are mentioned in the other question. Everything you mentioned is about the technology, but the real issue is the pilot: a new pilot simply can't handle how quickly things happen - and how quickly they can go wrong! - in a jet. As a student pilot, I was regularly 'behind the aircraft' in a C152, I have no idea how what would have happened in an aircraft moving three or four times faster.

"Operating a piston engine requires adjusting the mixture ratio, carburetor heating, ignition timing, and other such voodoo" Not necessarily. Certainly the Ikarus I'm flying is very forgiving in that manner, with little voodoo involved; electrics on, magnets on, choke if cold, press START, immediately verify oil pressure after engine start, and that's about it. Yes, it has selectable carb heating, but certainly outside of engine maintenance no need to worry about mixture or timing. Starting and flying that one is only slightly more involved than starting and driving a modern car.

It would be very easy to design a jet powered version of the Cessna 152. It could have all the same forgiving slow speed characteristics.The only reason it hasn’t been done is money. NO ONE would buy a $2,000,000 trainer that cruises at 90 knots.

There's also the concept of not letting the airplane get ahead of you.  Many pilots start training quite young - in the US, pilots can solo at 16 and be licensed at 17 in powered airplanes. That can be before they ge their drivers license even!   Going lower and slower lets a new pilot get used to keeping up with airplane that goes faster than a car, offers quite a different view and is navigated differently, and at a higher rate of speed (3-4 times faster than a car - landing and take off are faster too), can bring in radio play depending on where learning is occurring, etc.  And Weather!

@Mike Sowsun: And if you, as a flight school operator,  did have a $2 million jet trainer, how much would you have to charge students to pay for it? This article http://nation.time.com/2013/04/02/costly-flight-hours/ puts the hourly cost of a T-38 jet trainer at $9355/hr.  Not many people are going to pay that.

In a word, money.  Have you priced time in a turboprop or light jet, vs time in a C152 or other piston trainer?  Or actually buying one?  You can pick up a decent used single for about the cost of a new car, while turboprop or jet is likely to start at half a million, and go up from there.  And add maintenance costs to that.
Then there are a lot of pilots (I'm one) whose flying just doesn't include what a jet or turboprop is likely to be good at.  I like to fly into short strips, often grass or dirt (and sometimes no strip at all e.g. dry lake beds).  Is your jet or turboprop well suited for that?
Even for someone looking towards a career with the airlines, you can build a lot of the necessary time a lot cheaper in a single-engined prop plane.  For instance, one of my flight instructors started doing that, getting paid plus lots of experience dealing with the mistakes of student pilots.  Then he moved to a job flying a small cargo plane doing pickups for FedEx or someone like that, before getting enough hours to look at an airline job.


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Why did the B-52 never have a JT4A-powered version?
All versions of the B-52 use the Pratt & Whitney JT3C turbojet (except for the last version, the B-52H, which uses the JT3D low-bypass turbofan instead).
Problem: the JT3C is a horribly anemic engine, especially the early versions used on the A-E (producing a mere 44.5 kN per engine on the B-52A, increasing only to 46.7 kN/engine on the B-E models).  The engines used on the F and G did better, but still topped out at 61.2 kN each.  Due to this low engine thrust, the A through G models of the B-52 all required the use of water injection (with the associated weight and space penalties) to augment their engines' thrust for takeoff, without which the aircraft would have been incapable of becoming airborne.
For this reason, the other major users of the JT3C (the 707 and DC-8 airliners) promptly switched over to the more powerful JT4A turbojet when it (the JT4A) came on the market; the JT4A provided up to 77.8 kN of thrust, allowing the JT4A-engined versions of these aircraft to do away with the heavy and complicated water-injection system.
In contrast, the B-52 never switched over to the JT4A, even though, as a heavy bomber aircraft stuffed to the gills with fuel and ordnance, it would have benefited even more from the added engine power than the airliners did.
Why did the B-52 stick to the JT3C (until they finally jumped ship for the JT3D), and not go for the JT4A when it became available?

Ask the congress?

Another thing is, military aircraft is always designed for a specific task. If the task hasn't changed, then the aircraft doesn't need to change, either, e.g. the enemy didn't move further away geographically, so why increase the range? The opposite is true as well, i.e. if the original engine is insufficient, then how did the aircraft manage to enter service in the first place. Also military aircraft doesn't have high pressure of operation cost like commercial passenger or even freight flights so fuel cost won't drive the change, either.

Well, I can tell you that Boeing / the Air Force seriously considered re-engineing the B-52 with 4 turbofans.

@user3528438: Strategic bombers had the concept of "loitering" - unlike missiles, you could launch them and hold them in the air without attacking. This protects them against a first-strike attack. A re-engined B52 would be better at loitering


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Why was the DC-8-70 successful, but not the 707-700?
The first high-bypass turbofan small enough to use on a narrowbody jetliner was the GE/SNECMA CFM56, which entered service in the late 1970s as an upgrade engine for the KC-135 tanker; Boeing almost immediately began working on a CFM56-powered 707 (the 707 and KC-135 are very close siblings), the 707-700, with a prototype flying in 1979.  Boeing offered the 707-700 for sale in 1979 and 1980 (the last 707 version in their product catalog), but got not a single order, and the prototype 707-700 was eventually converted back to a 707-320C and fobbed off on the Moroccan Air Force.
In contrast, when a CFM56-powered version of the DC-8 became available in 1982 (the DC-8-70), the airlines jumped in with both feet, even though the DC-8 had been out of production for a decade (unlike the 707, which was produced all the way up until 1979, and would have lasted longer still had the 707-700 caught on), and would thus be, except for the engines, years behind where the 707-700 would have been.
Why did the CFM56 fare so much better on the DC-8 than on the 707?

I was a boom operator and then got a respectable job after the air force... I saw the RED B707-700 with CFM56's at Grissom AFB summer 1979....possible 1980. So quiet on approach and did a low-noise go around. Amazing

I think there is a bit of a misconception in your question.
The DC-8 was already out of production by the time the CFM56 was developed, so all of the DC-8 70 series are actually just retrofits of the 60 series aircraft already in use. So it’s really not correct to say “a CFM56-powered version of the DC-8 became available in 1982.” More like a CFM56 retrofit option for existing DC-8 60 series became available. They did not return to producing the DC-8. On the other hand, although the last civilian 707's were made in 1976, before the CFM56 was available, Boeing did manufacture new military versions of the 707 with CFM56 engines from 1986 to 1991.
As for why the civilian 707-700 was not pursued, the main reason is stated in the Wikipedia page:

Boeing abandoned the retrofit program, since it felt it would be a
  threat to the Boeing 757 program.

The 757 and 767 programs were already in the works when the CFM56 became available.
The 707-700 was really more of a test program for retrofitting the fleet of military 707 aircraft with the new engines. They didn't pursue a retrofit for the civilian fleet of 707's so they could sell more of the new 757 narrow body aircraft.
If you consider how many DC-8's were upgraded to 70 series (110) vs the number of KC-135’s retrofitted (500) plus 39 more new CFM56 equipped 707 variants, then the new HBP engine was much more successful on the 707.

It isn't an issue with the engines, but rather a question of why the DC-8 was more popular and common in commercial service, long after the 707.
You identified the key in your question: the KC-135 (-135s in general), and other variants like the E-3, E-6, E-8, C-18, and C-137. Because the US military had such a large fleet of the 707 family, they acquired as many of the spares, used parts and even used airframes, that became available and stockpiled them.
As a result, 707 became much more expensive to maintain for civil operators, so the economics of the plane became unfavorable versus the DC-8.

When in the test program, the reason given was the DC-8, series 60, had a longer life span than the commercial 707 aircraft existing.  The KC-135 had a lot less time on them. Worked on the program both at Boeing Field and Tulsa.


text:
Why couldn't the CRJ100/200 be equipped with triple-string autoland?
The gold standard for zero-visibility ILS approaches and landings is a triple-string autoland system, where the signals from the localiser and glideslope beams and the aircraft's radioaltimeter system1 are fed into the aircraft's autopilots and control its flightpath directly (with the pilots monitoring the system so that they can execute a go-around if the ILS signal is lost or an autopilot fails), cutting out the human middleman.  For redundancy, each of the aircraftside links in the chain (ILS receivers, autopilots, radioaltimeters, etc.) is present in triplicate, producing three parallel autoland strings, each constantly comparing its inputs and outputs to those of the other two strings, so that a failure in any of the components of the autoland system can be automatically detected and alerted to the pilots.
However, apparently, not all aircraft are capable of being equipped with a triple-string autoland, with the CRJ100/200 series being one of the major exceptions:

Simultaneously, pilot organizations globally were advocating the use of Head Up Display systems primarily from a safety viewpoint. Many operators in non-sophisticated environments without many ILS equipped runways were also looking for improvements. The net effect was pressure within the industry to find alternative ways to achieve low visibility operations, such as a "hybrid" system which used a relatively low reliability autoland system monitored by the pilots via a HUD. Alaska Airlines was a leader in this approach and undertook a lot of development work with Flight Dynamics and Boeing in this respect.
However, a major problem with this approach was that European authorities were very reluctant to certificate such schemes as they undermined the well proven concepts of "pure" autoland systems. This impasse was broken when British Airways became involved as a potential customer for Bombardier's Regional Jet, which could not accommodate a full Cat 3 autoland system, but would be required to operate in those conditions. By working with Alaska Airlines and Boeing, British Airways technical pilots were able to demonstrate that a hybrid concept was feasible, and although British Airways never eventually bought the regional jet, this was the breakthrough needed for international approval for such systems which meant that they could reach a global market.

Why couldn't the CRJ100/200 be equipped with a full triple-string ILS-III autoland system?  Even if it didn't normally come equipped with triplicate autopilots/ILS receivers/etc., those surely could have been retrofitted to the aircraft to give it full ILS-III autoland capability...

1: The localiser beam is used for lateral guidance throughout the entire approach until the aircraft taxis off the runway; vertical guidance is provided by the glideslope beam down to flare height, and the radioaltimeters thereafter.

I have experience on that program and can tell you that:
A:  While anything is possible, this would cost a fortune to implement, requiring a complete avionics overhaul.  For a single potential customer?  Snowball's chance in heck.  The customer would have had to commit to paying the development costs, many zillions, which would also have a snowball's chance in heck of happening.
B: By the mid 2000s, no development money was being allocated for that sort of product improvement on the 200. By then the program funding was generally only to deal with airworthiness issues and upgrades to deal with avionics obsolescence, as development money focused on the 700/900.
C: British Airways aside, there was no demand to speak of from the industry for that kind of capability, that anybody would be willing to pay for. Adding HUGs on the 700/900 to get Cat IIIa cert was more than sufficient.  On the 200, even to do that would have required an avionics suite upgrade that no regional operator would be willing to pay for (again, Snowball's chance in...).  Regional operators run on very thin margins.
In the US, by the late 2000s, Regionals would not even implement optional Service Bulletins unless there was a financial payback of ONE YEAR or less.  That meant that only SBs mandated by Airworthiness Directive were being incorporated.


text:
What does NSC mean in a METAR?
Here's a METAR:

SCEL 090000Z 15005KT 5000 HZ NSC 10/08 Q1020 NOSIG

What does NSC mean?

WMO Publication No. 306 - Manual On Codes - Volume I.1 - Part A, Section A, FM 15 METAR, FM 16 SPECI:

15.14.14 When no cloud below 1 500 metres (5 000 ft) or the highest minimum sector altitude, whichever is greater, and no cumulonimbus and no towering cumulus are forecast, and CAVOK is not appropriate, the abbreviation NSC shall be used.

Some countries have published different rules:

WMO Publication No. 306 - Manual On Codes - Volume II:

Netherlands

15.9.1.1 Abbreviation NSC: The abbreviation NSC shall not be used. As appropriate, clouds present at and above 5 000 ft also shall be reported.

New Zealand

15.9.1.1 Other than at Auckland (NZAA), Wellington (NZWN) and Christchurch (NZCH) NSC is not used, clouds are reported at all heights and SKC is used to report no cloud.

The Manual of Surface Weather Observation Standards for Canada does not have NSC in the METAR definition.
AIP USA, GEN 1.7 - Differences From ICAO Standards, Recommended Practices and Procedures, Annex 3 − Meteorological Service For International Air Navigation, Part II, Appendix III:

4.5.4 ... CAVOK and NSC are not used.

And the Federal Meteorological Handbook No. 1 also does not have NSC in the METAR definition.

NSC means No significant clouds.

Welcome to Av.SE.  While your answer is entirely believable, it would be better if you could link to a source to corroborate what you've posted.  I don't doubt your answer, but the general expectation here is that answers have some authority to them, not simply "somebody on the internet said it means {this}".

You can just Google it. https://www.skybrary.aero/index.php/Meteorological_Terminal_Air_Report_(METAR)


text:
Why does the 737 Original have an explicit hydraulic-quantity indicator for only the A hydraulic system?
The 737 (Original, Classic, NG, and MAX) has two primary hydraulic systems (A and B),1 which provide power to the aircraft’s flight controls and other aircraft components (like the wheelbrakes, landing-gear extension/retraction, and nosewheel steering).
On all 737s, the cockpit instrument panels have gauges which provide a readout of the pressure available in both of the primary hydraulic systems.2  Hydraulic quantity, however, is a different matter; on the 737 Original, the cockpit instrumentation features a quantity gauge only for the A hydraulic system, with the B system having to make do with a mere low-quantity warning light (the same type of indicator as with the standby hydraulic system, although on a different one of the instrument panels):

(Diagram originally by the National Transportation Safety Board, page 11 [page 29 of the PDF]; cropped and imageified by me.)
In contrast, the later 737 Classic has quantity gauges for each of the two primary hydraulic systems:

(Diagram originally by the National Transportation Safety Board, page 17 [page 41 of the PDF]; cropped and imageified by me.)
Why doesn’t the Original have quantity gauges for both main hydraulic systems, rather than just for the A system?

1: A third, standby system is unpressurised in normal flight, but can be activated by the flightcrew (or, in a few critical situations, automatically) in the event of a primary hydraulic failure; the 737 also has manual-reversion capability for the ailerons and elevators in the event of a complete hydraulic failure, one of only a few jetliners to still include this capability.
2: The standby hydraulic system, which is normally unpressurised anyways, merely has a warning light that comes on if the pressure within an active standby system falls too low.

You're still asking a lot of theoretical questions. I'd like to see questions from you like "how do I license a US-registered airplane in Korea" or "what's the best way to ferry a Cessna across the Pacific" or "can I replace my own alternator or must the FBO do it" or "how do I get this bolt off (pic)" or "how open are Russian authorities to fly-ins by South Koreans?"  :) :) :)

@Harper-ReinstateMonica: Bit hard to ask non-theoretical questions when I'm not a pilot (nor likely to ever be one) or an aircraft mechanic.


text:
What is the difference between a business jet and a regional jet?
At first glance, business jets and regional jets seem virtually identical - both are small jets, usually low-winged twinjets (unless they're by Dassault, in which case they're trijets) with tail-mounted engines and T-tails.
What differentiates, say, a Legacy 600 or a Falcon 9 or a G650 or a Learjet 75 from a DC-9 or a CRJ-100 or an F-28?  Is it merely who operates it?

At a design level, the biggest difference is in design for intensity of utilization.  That is, operating cycles per year.   An RJ designed as an airliner is intended to operate on average, about 2500 cycles per year.  A corporate jet typically operates under 500 cycles per year (although with a charter operator or fractional ownership operator or a corporate shuttle they may run 1000 cycles plus).
The utilization drives the level of robustness required in structural and system design. A purpose built corporate jet put into airline use would have much lower dispatch reliability because it's getting beaten to death (like putting a luxurious but fairly lightly built car into service as a taxi).  At the opposite end, you have regional airliners that are sold as business aircraft.  These airplanes are overbuilt for their role as corporate airplanes and have a reputation for being pretty bulletproof.
Production wise, the biggest difference is that corporate airplanes are built at a lower rate and are almost always delivered "green" to a completion center selected by the customer to get the interior done (by Supplemental Type Certificate), whereas regional jets get their bus interiors done in-house.  There are a few exceptions to this (some Global Express interiors are done in-house by the OEM).  And the interiors themselves are built to a completely insane quality standard - I mean, the interiors have to be absolutely perfect when someone has 5-10 million just into the interior.
On customer acceptance on a corporate jet, the customer's acceptance team will be going around the interior while the airplane cruises at altitude making sure all the panel gaps are still straight with the hull pressurized, in addition to making sure everything related to the interior works (like plumbing doesn't freeze etc).  The slightest flaw requires rework.  God help you if you ding the woodwork...

I would say its not quite who operates but how its operated. A business jet is typically outfitted to carry a small number of people in a very luxurious way while a regional jet is outfitted to carry a lot of people a short distance.
No one even says business jets have to be small or of a select list of airframes, you can order a 747 business jet if you have the cash. Similarly no one says a regional jet has to be small, we are now seeing 747SR's doing short haul flights in Japan.

From a legal standpoint, the difference may be in the operation type as well. Scheduled regional flights here in the US will occur under part 121 while charter business operations will generally occur under part 135 if you own the plane the operation may even occur under part 91. This mainly has differences on inspection intervals for the airframe and some other operational related things. Assuming you can land the plane on the runway and depart again safely these don't limit where you can go.


text:
Is flap-asymmetry protection on airliners still active even on the ground?
On aircraft with flaps, a general certification requirement is that it must not be possible for the aircraft to suffer a loss of lateral control secondary to an asymmetric-flap condition.  This requirement can be satisfied in one of two ways (many aircraft go for both):

Demonstrating that the aircraft remains controllable without requiring exceptional pilot strength or skill, even if the flap(s) on one side are fully up and those on the other side are fully down.
Including a flap-asymmetry protection mechanism, which automatically cuts off power (hydraulic or electrical, whichever the aircraft uses) to the flap-drive motors if the flaps on both sides fail to move in synchrony.

The benefits of flap-asymmetry protection in flight are obvious; however, in addition to their roles as high-lift devices in flight and high-drag devices during landing and rejected takeoff, flaps are also used to allow airliner passengers and cabin crew to safely egress through the overwing exits in the event of an emergency evacuation on land1 (low-slung aircraft, like the 737 and most regional jets, simply have the passengers slide down the flaps to the ground, while taller aircraft have inflatable slides that deploy over the lowered flaps).  In such a situation, as optimal use of the overwing exits depends on the flaps being fully down, having flap-asymmetry protection active would be counterproductive, as a jammed flap on one side would prevent a properly-functioning flap on the other side from providing a good evacuation path, and flap-asymmetry-induced lift asymmetry is harmless on the ground anyways (it produces a rolling moment to one side or the other, but an aircraft is physically incapable of rolling to either side when on the ground unless it’s already at or very close to flying speed, because the ground is in the way).
Is airliner flap-asymmetry protection inactive when on the ground (like many other systems protecting against things that are only a hazard when airborne, such as stall protection [stickshaker/stickpusher] and GPWS warnings) to prevent it from hindering an evacuation, or does it remain active even then?

1: In the case of a water evacuation, the buoyant characteristics of most airliners are such that the water generally comes all the way up to the wings, making the degree of flap deployment, or lack thereof, irrelevant.

I'm not aware of any transport airplanes that have demonstrated control authority with flaps full and zero at the same time.  You either have asymmetry detection (broken drive shaft detection) or you have a redundant drive line.  The Dash 8 1 thru 300s use a redundant drive flex shaft that spans the entire flap drive line to drive the shafts outboard of a broken main shaft.  They don't require asymmetry detection.  All of the Canadair and Embraer regionals use ends of drive system asymmetric movement detection that is always active.

It is active whenever the flaps are being moved by the hydraulic system on the 737; it isn't available when using the backup (electrical) system.
Allowing asymmetric flap extension before takeoff could have consequences as serious as asymmetric extension in flight. Plus, tying the protection to an air/ground sensor is one more way that the system might fail.
Evacuations are pretty rare events; the chance of having one along with a malfunction as described in the OP is extremely remote. Such an unlikely occurrence doesn't overcome the reasons to keep asymmetry protection available full-time.

The lack of electric asymmetry detection is probably because an asymmetry event (broken drive line) following a hydraulic drive system failure would be a double failure event and doesn't need to be accounted for in the failure effect risk analysis.

@JohnK  Yeah, that'd be (incredibly rare) squared. The QRH tells you to watch the needles (L and R flap position) during electric extension, so you HAVE asymmetry protection - just not automated protection. Plus, the electric motors move the flaps pretty slowly compared to the hydraulics, so any asymmetry that developed would have a slow enough onset the PF could notice it well before it became significant.

Are you just holding a spring loaded switch down to move them in that mode?

@JohnK Yes, the PM holds the switch & watches the flap position indicator. When you get to the desired position, or if an asymmetry starts to develop, releasing the switch stops the motor. The same switch pushed in the opposite direction can retract them (though not the leading edge devices - they're hydraulically actuated only), such as in the event of a go-around.


text:
What sort of control mechanisms (i.e., PID controllers) do modern FBW aircraft use?
It's a well-established fact that aircraft like the Boeing 777, and all Airbus passenger aircraft that were designed after the A300, feature a digital fly-by-wire control system. Considering the Boeing 777 as an example, the pilot flying applies force onto their yoke and one of the flight computers (I'm not aware of the architecture(s) of Boeing FBW systems) interprets the input and changes the control surfaces of the plane accordingly.
I am wondering what control paradigms are used by Boeing and Airbus to translate pilot settings (i.e. trim), accelerometer and gyroscope data from the IMU, and feedback from the aircraft's control surfaces to re-adjust said control surfaces.
Would a simple PID controller suffice? If so, how would a single PID system be able to manage feedback data from different sources (i.e. how could a PID system handle feedback from a gyroscope and Airbus side stick at the same time)?
I understand that the accuracy of PID systems are fully dependent on the feedback they receive, so I would assume that software methods of error correction like Kalman filter would be employed. However, how can computer hardware from the 80s (such as the Intel i386) be able to sample data quickly and compute Kalman filter within acceptable response times?
I appreciate all answers, and I apologize if my questions require analyzing proprietary technical documents.

I think computers are a lot faster than you seem to give them credit for.

It's also important to realized that there isn't a single CPU performing all of the monitoring and control loops, but rather 10's if not 100's (depending on the AC)

Kalman filters are not complicated and have been used for years in taking GPS data and filtering it at a rate of 1hz or slower. I386's are quite fast when you remove the overhead of a multi-tasking OS, and as the above comment says, multiple processors are used and are typically specialized for one duty so it requires a great many processors to run a modern aircraft.

@RonBeyer Kalman filters are not complicated if you neglect all numerical issues and assume that floating-point values are perfect substitutes for real numbers.


text:
How are modern fighters cleaned?
I found this photo of a fighter jet being cleaned, how does this process work?
(usatoday.com)

You squirt water at it and the water removes the dirt? That can't be what you're asking, because it's obvious. So what are you asking?

Related gallery: Wash rack at Eglin AFB. May be removing salt after flight over ocean.

The photo above may also be a demonstration of a chemical weapons purge in case of a CBR attack on a base or aircraft operating in the area.

The squirting part, obvious.  Less obvious: Are chemicals (surfactants, etc.) added to the water?  Do any precautions need to be taken to keep the water out of places it should not go?  Is the plane taxied through the wash, or is it powered off?  Are there different types of wash for different purposes?

@DavidRicherby obviously it’s about the procedure, not how water washes dirt away. Is the aircraft taxied through the spray or parked and powered down? Does the aircraft need prepping to prevent water getting into unwanted places etc.

The F-35 in the picture isn't being washed.  Fighters are still handwashed in a washrack.

Yes, they use a chemical to clean aircraft. Back when I was in the Air Force, we used PD-680 type II.
I'm pretty sure they don't use that any more though. PD-680 was discontinued (at least for most applications) around the turn of the century. Most applications that used to use PD-680 switched to PRF-680. I can't guarantee it, but I'd guess that's what they use to clean aircraft.

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


text:
What is the reason for this discrepancy regarding the CRJ100/200's stickpusher-activation angle?
The AC646 accident report contains a discrepancy regarding the angle of attack at which the stickpusher of the accident aircraft (a CRJ100, formally a CL-600-2B19, referred to in the aforelinked AAR as a CL-65) would have activated:

The stick pusher mechanism is designed to prevent the aircraft from entering an aerodynamic stall by applying a control column force to pitch the aircraft nose down as the aircraft reaches its computer-calculated pusher (stall) AOA. The Canadair Regional Jet Airplane Flight Manual (AFM) contains a chart (page 06-01-23) which allows calculation of the aircraft stick pusher (stall) speed for various aircraft weights. For 44 180 pounds, the aircraft’s estimated weight at the time of the accident, with the flaps set at 45 degrees and the landing gear down, the pusher activation speed would have been about 109 knots calibrated (109 knots indicated) at one g. At the moment of the aerodynamic stall, the aircraft was at a g loading of about 1.227, and the pusher activation speed would have been about 120 knots, the natural stall speed would have been about 116 knots, and the pusher activation AOA would have been about 13.5 degrees. [...] At the moment of stall, the flaps had retracted a few degrees from 45 and the g was increasing, both of which would cause a slightly higher pusher activation speed; however, there would be no effect on the pusher activation AOA.
Information from the flight recorders indicates the following events occurred during the attempted go-around: the stick shaker activated when the aircraft was at 129 knots as the pitch was being increased through 4 degrees; and the right roll and stall onset occurred and the warbler tone activated when the aircraft was at 124 knots and with the pitch at 9.7 degrees. At the time the warbler sounded, the left and right AOA vane readings were approximately 8.7 degrees and 9.4 degrees, respectively. The pusher did not activate because, while the right AOA reached its trip point, the left AOA did not. [TSB Aviation Occurrence Report A97H0011, page 8 (page 18 of the PDF), my emphasis.]

In the first of the paragraphs quoted, the stickpusher-activation angle of attack is stated as approximately 13.5 degrees (a value which, according to the report, would not have been altered by the aircraft's rapidly-increasing AoA); however, the very next paragraph implies that the stickpusher would have activated at an AoA of between 8.7 and 9.4 degrees (the right AoA sensor, reading 9.4 degrees, is stated as having reached the stickpusher point, while the left sensor, reading 8.7 degrees, had not reached the activation threshold by the time the aircraft stalled, thus implying that the threshold AoA for activating the stickpusher was above 8.7 degrees but below 9.4 degrees1), rather than at about 13.5 degrees.
How to reconcile this discrepancy?

1: Due, presumably, to the potentially-catastrophic results of a spurious stickpusher activation at low altitude, the CRJ100/200's2 stickpusher does not activate unless both of the aircraft's angle-of-attack vanes measure an AoA at least as high as the stickpusher-activation threshold; in contrast, all the automatic high-AoA protections that do not alter the aircraft's flightpath (continuous ignition, stickshaker activation, autopilot disengagement [same threshold as for the stickshaker], stall-warbler activation [same threshold as the stickpusher], and stall-warning-flashing-light activation [also same threshold as stickpusher]) activate even if just one of the two vanes measures an angle of attack equal to or greater than the requisite threshold.
2: Although the accident aircraft was a CRJ100, the CRJ200 is identical to the CRJ100 except for using a slightly-different variant of the CF34 engine; the two even share the same CL-600-2B19 classification (making it somewhat difficult, when reading AARs discussing first-generation CRJs, to figure out whether the aircraft being referred to is a CRJ100 or a CRJ200).


text:
How do airplanes act as corner reflectors despite lacking significant flat surfaces in one of the three necessary mutually-perpendicular planes?
One feature common to virtually all stealth aircraft is that they have vertical tails that are tilted away from actually being vertical (or lack them entirely).  This is said to be because the 90° angles formed by, on the one hand, the vertical tail of a typical aircraft, and, on the other, its wing or horizontal stabiliser, act as corner reflectors, reflecting incoming radar signals back towards their source and giving other people a nice bright return from the aircraft; having a non-vertical vertical tail, or no vertical tail at all, eliminates this effect, scattering radio waves away from their source and drastically reducing the aircraft’s radar cross-section.
However, a corner reflector, in order to function as such, needs three mutually-perpendicular flat surfaces (shaped like the inside of a corner of a hollow cube, hence the name) in order to reverse an incoming photon’s motion in the x, y, and z axes and send it back towards its source.  A reflector consisting of two perpendicular flat surfaces would reverse our photon’s motion in only two of the three spatial axes, sending it out on a track deflected 90° from its incoming trajectory, rather than the 180° necessary to return it to sender.
If we take an aircraft’s longitudinal axis (its roll axis, running from nose to tail) as its x axis, its vertical axis (its yaw axis, where you would stick a mounting pin through its fuselage1) as its y axis, and its lateral axis (its pitch axis, running from wingtip to wingtip2) as its z axis, aircraft, even non-stealth ones, generally only have flat surfaces in (or near) the xy (vertical tail[s], opened landing-gear and bomb-bay doors) and xz (wings, horizontal tails, canards) planes, or in other planes at least reasonably close to parallel with the aircraft’s x axis; aircraft tend to avoid having large flat surfaces in the yz plane, as these would create impressive amounts of drag.
So where does the corner-reflector effect of a typical airplane come from?

1: Think of the airplane as a giant butterfly being mounted in a display box.
2: For aircraft with highly-swept wings (displacing the wingtips forward or aft), or where the amount of lift produced by the secondary horizontal surface (horizontal tail or canard) is a large fraction of that produced by the main wing (requiring the wings to be displaced away from the aircraft’s center of mass in order to keep its overall center of lift close to the center of mass), the lateral axis may not actually pass through the wingtips, but this serves as a pretty good approximation.

I can't find hard numbers, but actual corner reflectors may reflect much more strongly than the typical airplane.  For instance, there's a 1957 patent for putting them in landing gear to "improve the reflective response" on takeoff or final.

I always understood that the radar reflective properties of a typical aircraft came from the curved surfaces (which offer mirror reflection from at least some points in any direction). That's why stealth aircraft have flat surfaces and sharp angles, so mirror reflection would tend not to return the beam to the sensor. (Disclaimer: Not a radar engineer.)

A 3-sided corner reflector will work for a beam coming from any direction on the right "side" of the corner. But a 2-sided reflector will still work for signals in the plane orthogonal to the two sides (i.e. in the YZ plane), like any surface will work for a signal coming on the line orthogonal to that plane. Don't think that's enough to explain things, though.

"… a corner reflector needs three mutually-perpendicular flat surfaces" - Yes, in 3D. Two surfaces suffice in one plane orthogonal to those two, and given several connected radar stations in the area it is easy to pick out the airplane from the frequent flashes it produces.

Generally airplanes have very little features that form  perfect trihedral corner reflectors. As Niels explained in his answer, such features are not necessary for a plane with metal parts to be clearly visible on a radar.
However, very small features that act as trihedral corner reflectors do so very efficiently. Such features might be engine mounts, engine intakes/exhausts etc. Radar reflectors are commony used in sailing vessels, hence the link to the following article: Selecting a Radar Reflector
, and a quote from it:

The effectiveness of an RTE is disproportionately related to its size. Assume that you have three theoretical reflectors of the same design, but of different sizes. Look at how rapidly the RCS (Radar Cross Section) increases with size. The RCS of a given reflector goes up by the fourth power of the radius, resulting in this dramatic increase in effectiveness.

The article continues to give following theoretical examples of RCS for certain radii of RTE's:

3" = 1.0m2
5" = 7.5m2
7" = 28.8m2

(Really don't understand why they mix imperial and metric units, but well...)
So: even a very small feature either acting as a trihedral corner reflector, or a dihedral corner reflector at a suitable agle will return a disproportionally large radar echo. Keep in mind that conventional aircraft fuselages have large areas where two curved surfaces blend, mainly the wing-body interface. Such areas create a sort of a three dimensional dihedral corner reflector that sweeps a large agle across space. These corner reflectors, though not trihedral, are very "visible" to radars.
The above is quite obviously the reason any trihedral features are avoided in stealth aircraft.
Now: it is unpractically hard to completely avoid features that act as dihedral corner reflectors in stealth aircraft. One such example is the F-117 Nighthawk, an ugly sob... Another, more aesthetic example is the B-2 bomber, which hardly is very maneuverable due to lack of efficient flight control surfaces. Modern stealth aircraft such as F-22 and F-35 accept some possible dihedral corner reflection spots, but mitigate the effect by use of non reflective / absorptive materials.
Epilogue: Modern civil radars do not need much of a return signal to spot a target, military radars even less so. The internet provides a plethora of comparisons (such as this) of RCS's of aircraft, generally ranging from the equivalent of an insect for F-22 and B-2, to the size of a car for the good old B-52. Be adviced that some modern military radar systems are able to target and track even the stealth aircraft from substantial distances (source will not be provided, dwi :)

most excellent, +1- NN

A corner reflector is the most efficient way to generate a strong radar return from an incoming pulse. However, even if the plane has no corner reflectors in its overall structure, it will still generate a return because it is made of metal, and the characteristic impedance of the metal is sufficiently different from that of air to generate a reflection of the incoming electromagnetic wave. That reflection will not be collectively "aimed" straight back to the source is in the case of a corner reflector but if the incoming wave is strong enough and the radar receiver sensitive enough, the plane will still show up on the radar screen.

Of course. The question asked How come there is a corner reflector on the plane

To be excact, it does not.

Even if the surfaces aren't exactly orthogonal to each other, the horizontal and vertical tails produce good enough corner reflectors to make the airplane much more visible in side view. Add rectangular intakes with ramps for supersonic flight and they really shine up. Those spots regularly require angles of less than 70° or more than 110° to lose their corner reflectivity.


text:
Does the MD-90 use a different elevator system from the other DC-9 versions?
While reading the 7Z9363 (a 2017 DC-9-80 [“MD-80”] very-high-speed rejected takeoff and runway overrun resulting from an undetectable wind-induced elevator jam that rendered the aircraft unable to rotate for takeoff) accident report, I noticed a curious and persistent omission (my emphasis):

...According to Boeing, in the history of this elevator design (which exists on all Boeing DC-9/MD-80 series and 717 model airplanes), this accident was the first notification that Boeing had received of an elevator jam occurring on an airplane exposed to ground gusts lower than 65 kts. Boeing noted that the elevator design first entered service in 1965 on the then-Douglas DC-9 airplane.  [NTSB AAR-19/01, page 12 (page 28 of the PDF).]
In November 2018, Boeing provided an update that included copies of revised Fleet Team Digest (FTD) documents that it had distributed to operators... According to the FTDs, Boeing was developing a service bulletin (SB) for modifying the elevator structure of Boeing DC-9/MD-80 series and 717 model airplanes to attach a secondary travel stop that would prevent the excessive elevator TED travel that could result in the geared tab linkage becoming locked overcenter.
Boeing also reported that it was developing a revision to the AMM for Boeing DC-9/MD-80 series and 717 model airplanes to add new elevator wind damage inspection procedures, which would also include a lower wind speed threshold for the inspection. The inspection would involve verifying that an elevator was positioned trailing edge neutral or above.  [Page 48 (64).]
The following analysis discusses the flight crew’s performance, including their preflight airplane inspection and control checks, preflight weather planning, and execution of the rejected takeoff after V1 (section 2.2), and evaluates the following:
[...]

Mitigations to prevent takeoffs with an undetected jammed elevator for Boeing DC-9/MD-80 series and 717 model airplanes (section 2.4);  [Page 50 (66).]

The NTSB notes that, given the circumstances of this accident and the potentially catastrophic outcome of an undetected jammed elevator, effective mitigations are needed. Therefore, the NTSB recommends that Boeing complete the development of a modification for Boeing DC-9/MD-80 series and 717 model airplanes that will prevent the possibility of elevator jamming due to ground wind exposure.  [Page 58 (74).]
Thus, the NTSB concludes that, to ensure that the elevators of Boeing DC-9/MD-80 series and 717 model airplanes are inspected by maintenance personnel when exposed to ground gusts that meet or exceed criteria specified in the AMM, operators of these airplanes must maintain awareness at all times of the forecast and known wind where the airplanes are parked.  [Page 59 (75).]

Missing from almost all of this discussion of DC-9 elevator jams is any mention of the MD-90, the largest DC-9 ever built, the last to be produced by McDonnell Douglas,1 and the penultimate DC-9 version overall.  The report mentions the MD-90 only once, in connection with a 2001 Boeing operational bulletin (issued as a result of a 1999 incident) regarding wind-induced elevator jams:

Subsequently, on June 25, 2001, Boeing issued Flight Operations Bulletin (FOB) MD-80-01-02, “Flight Control Jam,” that alerted operators about the circumstances of the incident, including the belief that the wind forced the left elevator into a TED position beyond the design limits, causing it to jam.35
[...]
35The FOB, which applied to all Boeing DC-9, C9, MD-80, MD-90, and 717 airplanes, was also referenced by number DC-9-01-02, MD-90-01-02, and B-717-01-05.  [AAR-19/01, pages 40-41 (56-57).]

This seems to indicate that wind-induced elevator jams were a concern for the MD-90 in 2001 (when the original bulletin regarding said jams was issued, applicable to all versions of the DC-9 - classic, -80, -90, and 717), but not in 2017-19, at the time of the 7Z9363 accident and aftermath thereof (all the discussion of said accident, and of the actions taken following same, mentions the classic DC-9, the -80, and the 717, but not the -90).  Was the MD-90 given a different elevator system, or its preexisting elevator system modified, between 2001 and 2017, in a manner not true for any other version of the DC-9?

1: Assembly of the first prototype of the MD-95, the very last DC-9 version, started under McDonnell Douglas ownership, but McDonnell Douglas had merged with Boeing by the time it was rolled out bearing the 717-200 moniker.


text:
Why are end-fire glideslopes less useable for low-minimum approaches than image-type glideslopes?
In this answer to an earlier question of mine regarding the predominance of image-type glideslopes (which bounce the lower glideslope beam off the ground rather than transmitting it directly up and out) over end-fire glideslopes (which use a separate antenna installation for the lower beam and don't need to bounce it off the ground), @UnrecognizedFallingObject pointed out something surprising about the differences between the capabilities of the two types of glideslope:

Note also that there are multiple flavors of image-type glideslope antenna (null reference, sideband reference, and capture effect), and some are more tolerant of poor site conditions than others. Furthermore, even though endfire glideslopes can work in situations where an image-type glideslope cannot, they have their own limitations as well. For instance, while well-sited image-type glideslopes can provide the tight tolerances required for a CAT IIIb ILS, end-fire antennas are limited to CAT II operations under ideal conditions, and systems designed for the most adverse sitings are apparently limited to CAT I usage. [Emphasis added.]

Given that both an image-type and an end-fire glideslope transmit the upper glideslope beam in the same manner, and that an end-fire glideslope transmits the lower beam directly up and back along the approach path, while an image-type glideslope has to bounce it off the ground first, it seems that the end-fire glideslope should have the easier time avoiding distortion or interference, as the end-fire glideslope's lower beam

is not reliant on reflection from the ground, eliminating the potential for beam distortion from irregularities in the ground surface, and
spends much less time in close proximity to the ground (as it starts moving upwards as soon as it leaves the antenna, whereas an image-type glideslope's lower beam is transmitted downwards, descends to ground level, bounces, and only then starts moving upwards), greatly reducing the potential for interference from terrestrial sources like airport vehicles or local wildlife,

which would, at first glance, indicate that the end-fire glideslope should be more able to provide tight ILS-III tolerances than is the image-type glideslope.
Yet, apparently the opposite is the case (and, frustratingly, the FAA document linked in the abovequoted answer makes no mention of this problem) - why?

NAA since I don't have anything but my own theorizing to go on here, but there's a possibility that end-fire systems simply haven't been tested to lower minima for some reason or another


text:
Why would larger aircraft have higher crosswind takeoff and landing limits than smaller aircraft?
In a response to an earlier question about the paucity of runways at major British airports, a major theme was that the advent of larger, heavier aircraft with higher maximum-crosswind limits for takeoff and landing (combined with Heathrow’s relatively consistent windfield) reduced or eliminated the need for crosswind runways:

When Heathrow was opened as a commercial site it had THREE different runways in a triangle and by 1955 it had SIX runways - You can see them here: Wikipedia [sic] commons - arranged to allow parallel operation on any 2 runways no matter what the wind direction was.
But with the coming of larger transport aircraft having higher landing speeds and greater crosswind tolerances, the need for the extra runways was diminished - and by the end of the 1950s only the east/west runways were being used - they got extended into the 2 runways in use today, whilst the other runways were closed - they're used as taxiways today.

However, the things determining how much takeoff or landing crosswind an aircraft can handle are:

First and foremost, the aircraft’s maximum rudder authority, which, most prominently, determines how rapidly the pilot can decrab the aircraft upon touchdown during a crosswind landing, in order to keep the main-gear tyres (which, for most aircraft,1 are locked in a fore-and-aft orientation and cannot caster to align themselves with the aircraft’s direction of motion along the ground, unlike the nose gear) from being destroyed by the sideways forces generated by their being dragged down the runway askew; it also determines how much crosswind the aircraft can tolerate during its takeoff or landing roll before it starts weathervaning uncontrollably into the wind due to the force of the wind on the aircraft’s vertical stabiliser (although this is usually a secondary concern, as it generally requires less rudder authority than the large, high-rate yaw manoeuvre required to decrab the aircraft upon touchdown in a strong crosswind).

The aircraft’s maximum rudder authority, in turn, is determined by the ratio between, on the one hand, the amount of yawing torque the rudder is capable of exerting on the aircraft (a function of the rudder’s size, its effective moment arm, and [to a lesser degree] its maximum deflection angle) plus the additional yawing torque produced by the forward fuselage with the aircraft at a nonzero sideslip angle (a function of its size and effective moment arm, and of the aircraft’s sideslip angle), and, on the other, the amount of weathervaning torque produced by the vertical stabiliser and aft fuselage when the aircraft has a nonzero sideslip angle (a function of the size and effective moment arm of the aft portions of the aircraft, and of the aircraft’s sideslip angle); as all of the areas and effective moment arms involved would be increased or decreased in essentially equal proportion by a change in the aircraft’s overall size, a larger aircraft would not be expected to have significantly greater rudder authority than a smaller one.
A much more prominent factor in the rudder-authority department is, for multiengine aircraft with non-centerline thrust, the need to use the rudder to maintain directional control in the event of an engine failure during takeoff, with the good engine still at full power, at speeds down to far below those required to maintain flight;2 the minimum airspeed at which the rudder has enough control authority to do so is referred to as VMC (minimum controllable speed), and forms a hard lower limit on the aircraft’s range of allowable V1 speeds (and, thereby, also on the amount of runway it requires for takeoff at a given weight).  For non-centerline-thrust multiengine aircraft, this, rather than crosswind landings and takeoffs, is generally the critical case that defines the amount of rudder authority needed; thus, if there were to be a dichotomy in aircraft’s maximum rudder authorities, one would expect it to be between single-engine (and centerline-thrust multiengine) aircraft, on the one hand, and non-centerline-thrust multiengine aircraft, on the other, rather than between small aircraft and large aircraft per se.3  If anything, one would expect some small aircraft with wing-mounted engines to need more maximum rudder authority (and, thus, have greater crosswind capabilities) than some large aircraft with engines mounted close to the aircraft’s centerline!

Secondly, the ability of the aircraft’s main-gear tyres to withstand the considerable sideways forces exerted on them before the aircraft is fully decrabbed; this determines the maximum crab angle at which the aircraft can safely touch down without tearing the tyres off its main landing gear, as well as the amount of time available for decrabbing the aircraft from a given angle before serious tyre damage occurs.  If anything, smaller aircraft would seem to have an advantage here, as...

Smaller aircraft tend to land and take off at lower airspeeds, thereby minimising the sideways force on the main-gear tyres for a given crab angle, and, thus, presumably, increasing both the maximum safe touchdown crab angle and the safe time available for decrabbing from a given angle.
Smaller aircraft tend to have a lower tyre loading (less weight per unit tyre-contact-patch area) than larger aircraft, reducing the frictional force between the tyre tread and the runway surface, and, thus, the sideways forces on the main-gear tyres at a given airspeed and crab angle.

So why would larger aircraft have greater crosswind-takeoff-and-landing limits than smaller aircraft, rather than the other way around?

1: Although not all.
2: It is also generally considered desirable to maintain control of one’s aircraft in the event of an engine failure during cruise flight or when landing; however, these are much less critical situations in this regard, as...

During cruise, an aircraft’s airspeed is generally much higher than during takeoff or landing (greatly increasing the rudder’s control authority), and its engines are at the considerably-lower-power cruise thrust setting (reducing the maximum thrust asymmetry - and, thus, the maximum rudder authority needed - should one fail en route), rather than the maximum-power TOGA (TakeOff/Go-Around) setting generally used during takeoff.  (Some takeoffs are made light enough, and from long enough runways and/or into strong enough headwinds, to allow one to safely use a lower-power “flex” setting for takeoff, which decreases engine wear, fuel consumption, and noise production compared to using full TOGA power, at the expense of lengthening the aircraft’s takeoff roll and lowering the maximum allowable takeoff weight; however, even this setting still uses a much higher thrust level than the cruise setting.)
During landing, the aircraft’s engines are generally spooled down almost to flight idle for the descent, greatly reducing the maximum thrust asymmetry (and, thus, maximum rudder authority required) in the event of an engine failure; even if TOGA power is needed (for instance, during a go-around or an encounter with strong windshear), the aircraft’s airspeed is still significantly higher than it is during the most critical portions of the takeoff ground roll (and the directional-control margins accordingly greater).

3: Granted, nearly all single-engine aircraft are also quire small... but so are many, many non-centerline-thrust multiengine aircraft, so it wouldn’t be just large aircraft having greater crosswind capabilities - many small aircraft would be expected to as well.

The overall trend has to do with speed. A rough check on the crosswind capability of an aircraft at a particular flap configuration is its sideslip capability, which roughly corresponds to the condition of a fully de-crabed crosswind landing. With a similar vertical tail volume and rudder chord ratio, two airplanes of very different sizes will have fairly similar sideslip capability.
Let's do a quick check:

Cessna 172 has a final landing speed of approximately 61kt. It has a maximum demonstrated crosswind of 15kt. This corresponds to a sideslip of 14deg.
B737-500 has a MZFW of 45 metric tons and a corresponding Vref of 122kt at Flap 40. It has a max demonstrated crosswind of 35kt. This corresponds to a sideslip of 16deg.

Disclaimer: do not use the above figures for flight. They are for references only.
Of course, a successfully demonstrated crosswind landing does not need to be fully de-crabed; landing gears have margin built in to cater for a crabbed landing and the associated lateral loads. Nevertheless, sideslip capability is a good first indication of a plane's crosswind capability. It's so important that steady-heading sideslip maneuvers are performed in development flight testing to ascertain stability at full pedal input prior to demonstrating maximum target crosswind.

Good answer - I'd also add that other elements of aircraft design also factor in. For example, a KC-135 has a wing mounted under the fuselage with upgraded engines mounted under the wings, leading to reduced engine-ground clearance.  The restricted bank angle from low mounted engines on touchdown limits the ability to come out of a sideslip and land safely,  thus there are crosswind limits not capped by aerodynamic factors.

You generally don't land heavies using cross controlled sideslip. The method is flare in the crab wings level, yaw the nose to align with the runway, and plant it before any significant drift starts in the direction of the skidding turn you've initiated by yawing wings level.  The inertia of the mass gives you time to complete this maneuver without having to lower the wing as long as you don't delay the touchdown.

@JohnK That's the exact maneuver we do crosswind demonstration with. We don't hold any slip prior to flare.

I don't agree with this at all, the C172 has half the crosswind limit of a Bulldog, but both have almost identical speeds for landing.

@GdD Good point. Bulldog has a much smaller tail length than the 172 at much higher wing loading. From pictures, it seems that Bulldog also has more rudder chord than 172. Makes sense that it would have more yaw capability at the expense of reduced yaw stability.